SACCON Forced Oscillation Tests at DNW-NWB and
NASA Langley 14x22-foot Tunnel
Dan D. Vicroy*
NASA Langley Research Center, Hampton, VA 23681
Thomas D. Loeser^
German-Dutch Wind Tunnels, Braunschweig, Germany
and
Andreas Schutte J
DLR Institute of Aerodynamics and Flow Technology, Braunschweig, Germany
A series of three wind tunnel static and forced oscillation tests were conducted on a
generic unmanned combat air vehicle (UCAV) geometry. These tests are part of an
international research effort to assess the state-of-the-art of computational fluid dynamics
(CFD) methods to predict the static and dynamic stability and control characteristics. The
experimental dataset includes not only force and moment time histories but surface pressure
and off body particle image velocimetry measurements as well. The extent of the data
precludes a full examination within the scope of this paper. This paper provides some
examples of the dynamic force and moment data available as well as some of the observed
trends.
A
b
C A
C N
C Y
Q
C m
r
C r
Gef
P
q
r
t
amplitude of oscillation
Nomenclature
V
_
velocity
span
a
=
AoA, angle of attack
axial force coefficient
a
=
angle of attack rate
normal force coefficient
P
=
sideslip angle
side-force coefficient
P
=
sideslip angle rate
rolling moment coefficient
V
=
yaw angle
pitching moment coefficient
0
=
pitch angle
yawing moment coefficient
9
=
roll angle
root chord
CO
=
angular velocity
reference chord
ESP
=
electronically scanned pressure
oscillation frequency
MRP
=
moment reference point
reduced frequency
PIV
=
Particle Image Velocimetry
roll rate
RLE
=
round leading edge
pitch rate
RLE-FT
=
RLE with fixed transition
yaw rate
SACCON
=
Stability and Control Configuration
time
SLE
=
sharp leading edge
Senior Research Engineer, Flight Dynamics Branch, MS 308, NASA LaRC, AIAA Associate Fellow.
1 Project Manager, DNW-NWB, Filienthalplatz 7, 38108 Braunschweig, Germany.
* Research Engineer, DER Institute of Aerodynamics and Flow Technology, Lilienthalplatz 7, 38108 Braunschweig,
Germany.
1
American Institute of Aeronautics and Astronautics
I. Introduction
T HE role of computational fluid dynamics (CFD) methods in the design process of both air and sea vehicles
continues to grow. However, the ability of CFD to accurately predict the static and dynamic stability
characteristics of these vehicles has yet to be validated. An Applied Vehicle Technology Task Group (AVT-161)
was established by the NATO Research and Technology Organization (RTO) to assess the state-of-the-art in CFD
methods for the prediction of static and dynamic stability and control characteristics of military vehicles in the air
and sea domains 1,2 . Two highly swept wing configurations were selected for the air vehicle experimental and
numerical investigations. The primary configuration was a generic UCAV geometry called SACCON - “Stability
And Control CONfiguration”. The SACCON model mounted in the Low Speed Wind Tunnel Braunschweig
(DNW-NWB) is shown in Fig. 1. The other focus air vehicle of AVT-161 was the X-31 3 " 5 .
This paper presents the results of a series of forced
oscillation tests of the SACCON model conducted in
both the DNW-NWB low speed wind tunnel and the
NASA Langley 14-by 22-Foot Subsonic Tunnel. A
complementary set of static data was also collected
during these tests and reported in reference 6.
II. Test Setup
A. SACCON Model
The SACCON UCAV has a lambda wing planform
with a leading edge sweep angle of 53°, see Figs.
2 and 3. The root chord is approximately lm and the
wing span is 1.53 m. The main sections of the model are
the fuselage, the wing section and wing tip. The
configuration is defined by three different profiles at the
root section of the fuselage, two sections with the same
profile at the inner wing, forming the transition from the
fuselage to wing and the outer wing section. Finally the
outer wing section profile has 5° of washout about the leading edge to reduce the aerodynamic loads and shift the
onset of flow separation to higher angles of attack.
The leading edge was the only exchangeable part of the model, providing a sharp (SLE) and a variable round
leading edge (RLE). The RLE is sharp at the root chord and the leading edge radius is growing in the span-wise
direction up to the intersection between fuselage and wing and then decreasing again.
Figure 1. SACCON low speed wind tunnel model on
the MPM-“Model Positioning Mechanism” in the
closed test section of the Low Speed Wind Tunnel
Braunschweig (DNW-NWB).
Figure 2. Planform and geometric parameters of the
SACCON UCAV configuration.
Figure 3. SACCON UCAV configuration with force,
moment and angle orientations in body and wind
axes.
2
American Institute of Aeronautics and Astronautics
Figure 4. Pressure tap location on the upper surface
of the SACCON configuration (Surface pressure
contour from preliminary CFD calculation).
The model is made of carbon fiber reinforced plastic Fi S ure 5 - Pressure tubes and ESP modules of the SACCON
and is very light with an overall weight of less than 10kg w ' nd tunne * modeb
(including pressure tubes and ESP modules). The very light design reduces the dynamic inertial loads enabling the
use of a smaller, more sensitive balance that provides better force and moment resolution.
The SACCON wind tunnel model is equipped with more than 200 pressure taps on the upper and lower side of
the model. The taps are connected with pressure tubes to electronically scanned pressure (ESP) modules within the
model. At ten additional positions unsteady pressure transducers are mounted. The location of the pressure taps and
transducers are depicted in Fig. 4. The pressure tap locations were selected based on preliminary CFD computations.
The aim was to capture the complex vortex flow topology over the configuration at operation points of the
trajectory. All pressure tube connections between the pressure taps and ESP modules are of the same length to
guarantee the same time dependent behavior for each pressure tap during the unsteady pressure measurements. This
leads to big bundles of the flexible tubing which have to be carefully installed to prevent kinks. The tubes bundles
which have to be placed inside the model are shown in Fig. 5.
Initial tests with the RLE configuration showed a variable transition line on the upper surface of the model
detected by infrared thermography 6 . These measurements led to the decision to prepare the leading edge with a
carborundum grit trip as it is shown in Fig. 6. The grit
was applied to approximately the first 25 mm at the nose
to 10mm at the wing tip along both the upper and lower
surface. Subsequent infrared thermography showed that
after establishing the grit a fully turbulent flow over the
upper wing surface could be assumed throughout the
oscillation cycle. However, there were no dynamic runs
with replicated test conditions with which to compare the
effect of the grit on the dynamic forces and moments. All
of the round leading edge data presented in this report
will be with the carborundum grit. Conversely, all of the
sharp leading edge (SLE) data is without grit.
B. Wind Tunnels
The dynamic data for the SACCON model was
collected over three wind tunnel test entries in two tunnel
facilities. The first two tests were conducted in the
DNW-NWB Low-Speed Tunnel located on the DLR site
in Braunschweig, Germany, shown in Fig. 7. The last test
was conducted in the NASA Langley 14-by-22-Foot
Subsonic Tunnel in Hampton, Virginia, USA, show in
Fig. 8. Both of these wind tunnels are closed-circuit,
atmospheric facilities that can be operated with open or
closed test sections. The DNW-NWB test section size is
w
Figure 6. Leading edge with carborundum grit trip
on the RLE-FT configuration (FT: fixed transition).
3
American Institute of Aeronautics and Astronautics
Figure 7. DNW-NWB low-speed tunnel in
Braunschweig, Germany.
Figure 8. NASA Langley 14-by-22-Foot Tunnel in
Hampton, Virginia, USA.
3.25 m by 2.8 m (10.6' by 9.2'). The maximum free stream velocity is V=80 m/s (263 ft/s) with the closed test
section and V=70 m/s (230 ft/s) in the open test section. The test section of the Langley 14-by-22-Foot tunnel is
4.42 m by 6.63 m (14.5' by 21.75') with a maximum free stream velocity of V=106m/s (348 ft/s)
provides an illustration of the relative size of the tunnel test sections
with the SACCON model in its mounting orientation.
Figure 9
Figure 9. Illustration of SACCON model
mounting orientation and relative test
section size.
C. Forced Motion Systems
1. MPM system
DNW-NWB’ s Model Positioning Mechanism (MPM) is a six
degree-of-freedom (DOF) parallel kinematics system designed for
static as well as for dynamic model support. Characteristic features
of this unique test rig are the six constant length struts of ultra high
modulus carbon fiber and the six electric linear motors, which move
along two parallel rails. The first Eigen frequency at the MRP is
above 20 Hz. The MPM is located above the test section and can be
operated in the open test section as well as in the closed one. The
location of oscillation axes can be chosen arbitrarily and in addition to classic sinusoidal oscillations the MPM can
perform multi-DOF maneuvers. The model location and orientation in the tunnel are determined through an optical
photogrammetric system featuring two high speed video cameras. The cameras have been mounted below the test
section and acquire 1280x1024 pixel images at 300 frames per second, each. The position and attitude of the model
are calculated in real time from the pixel coordinates of three markers, which have been applied to the model
surface.
An artist’s impression of the MPM carrying the SACCON model is shown in Fig. 10. Although the illustration is
depicted with an open test section the SACCON test was
conducted in closed configuration. More details concerning
the MPM are given in Bergmann et al. 7,9
Although the MPM system is capable of complex
simultaneous multi-axes motions the SACCON model was
only tested with single axis, constant amplitude and
frequency, sinusoidal motions. The MPM system was used
for pitch and yaw oscillations with ±5° amplitude at
frequencies from 1 to 3 Hz. Plunging oscillations were also
conducted with ±5 0mm amplitude at frequencies of 1 and
2.5 Hz.
2. NASA Forced Oscillation Rig
The forced oscillation (FO) test rig in the NASA Langley
14-by-22-Foot Subsonic Tunnel (shown in Fig. 11), can
provide constant amplitude and frequency sinusoidal motion
in the roll, yaw or pitch axes 10 . The frequency can be set
from 0.05 to 1.0Hz at amplitudes up to 30 degrees. The Fi S ure 10 ‘ SACCON on the MPM support in the
F F & DNW-NWB wind tunnel.
4
American Institute of Aeronautics and Astronautics
Figure 11. SACCON model mounted on the forced Figure 12. SACCON model with rear sting mount,
oscillation test rig for yaw oscillations in the NASA
Langley 14-by-22-Foot Subsonic Tunnel.
model angle of attack is set my rotating the turntable on which the FO rig is mounted. The oscillation angle of the
model is measured from an angular position transducer on the FO rig. The NASA rig was used to test the SACCON
model in the roll and yaw axes at oscillation amplitudes of ±5°, ±10° and ±15°. Some of the yaw oscillation tests
repeated conditions tested in the DNW-NWB Low Speed Tunnel for tunnel to tunnel comparison.
D. Mounting arrangements
The model was designed to accommodate mounting with a rear sting for oscillating in roll, as shown in Fig. 12,
or a belly sting for oscillating along the pitch or yaw axis, as shown in Figs. 11 and 13. Different connection links
between belly sting support and internal balance at DNW-NWB provide an angle of attack range from -15° to 30°.
This is provided by two different rigid cranked yaw links or by using an internal pitch link driven by a 7 th axis. The
two different basic setups with and without the 7 th axis are shown in Fig. 13. The location of the belly sting
connection well aft of the moment reference point (MRP)
was chosen to minimize the influence of the sting on the
overall flow topology. Previous investigations with the X-3 1
configuration have shown that for the prediction of the total
forces and moments the sting support has to be taken into
account 11 .
It should be noted that since the belly sting attachment
and subsequent rotation axis are well aft of the MRP or likely
SACCON center of gravity (eg) location the pitch and yaw
oscillation data from these tests are not representative of the
SACCON dynamic response when rotated about the eg. This
data is however valid for the intended purpose of comparing
with CFD predictions.
It can be seen in Fig. 13 that the connection between the
sting support and internal balance is completely covered by
the model fuselage for the configurations with yaw link.
These are adapted new designs especially for the SACCON
configuration. For the pitch link it was not possible to adapt
the design and a cover was used to smooth the geometry in
this area.
E. Forces and Moments
An internal six-component strain gauge balance was used
for the force and moment measurements. The DNW-NWB Figure 13. Top: 15° cranked yaw link support,
wind tunnel test used an Emmen 196-6 balance, whereas the Button: Support with 7 th axis and internal pitch
NASA 14x22-foot test used a FF-10D balance. These link.
5
American Institute of Aeronautics and Astronautics
balances were selected based on comparable load ranges and desired accuracy.
Note that the rolling and yawing moment coefficients were computed using the semi-span rather than span as the
reference length. The reference length for the pitching moment coefficient was the reference chord.
F. PIV Measurements
Particle Image Velocimetry (PIV) measurements were performed with two independent systems simultaneously
during the second DNW-NWB test by teams from ONERA and DLR. The majority of the PIV measurements were
taken statically, however measurements with the model oscillating in pitch were also performed. The PIV cameras
were triggered in a phase locked mode with the model oscillation. PIV data were recorded at eight different phase
angles over the oscillation cycle. Details of the PIV static and dynamic measurements are provided in references 15
and 16.
G. Test Matrix
The test matrix of model leading edge configurations, mounting arrangements, tunnel conditions and motion
parameters for each of the oscillation axes are listed in the appendix in Tables 1 through 4 and illustrated in Figs. 14-
lb. The motion parameters that were varied were the frequency and amplitude of the sinusoidal oscillation along
with the tunnel velocity. The corresponding reduced frequency ( k ) or Strouhal number and reduced angular rate are
shown in the Figs. 14-16 and Tables 1-4. The reduced angular rate is proportional to the tangent of the helix angle of
the vehicle rotation relative to the free-stream velocity. For example, the reduced angular rate in roll is proportional
to the induced angle of attack at the wing tip due to the rolling motion. The roll angle (cp) during the sinusoidal
oscillation is:
cp{t) = cp 0 + A sin (cot)
The roll rate is:
( p(t ) = p = Ao) cos (cot)
The maximum roll rate is:
Pmax =Aco = 2nfA
( 1 )
( 2 )
( 3 )
The reduced rotation rate and frequency are non-dimensionalized by the reference length divided by twice the free-
stream velocity, with the reference length being the span for roll and yaw and the reference chord for pitch.
0 10 20 30 40 50 60 70 0.00 0.02 0.04 0.06 0.08 0.10
V, m/s
Pmaxb/2V
Figure 14. Roll oscillation frequency, velocity and amplitude test points.
6
American Institute of Aeronautics and Astronautics
0.50
f, Hz
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J*
20 30 40 50 60 70 0.00
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Figure 15. Yaw oscillation frequency, velocity and amplitude test points,
0.02 0.04 0.06 0.08
'maxb
Pmax ~
Pmaxb
2V
, cjb
k = —
2V
( 4 )
( 5 )
Figure 14 shows that for the roll oscillations the frequency effect was explored with the round leading edge with
fixed transition (RLE-FT) configuration at tunnel velocities of 18, 35 and 43 m/s. Similarly, the effect of velocity
was tested at reduced frequencies of 0.097 and 0.064. The rate or amplitude effect was also tested at several reduced
frequencies with both the RLE-FT and SLE configurations.
Fewer frequency and amplitude combinations were tested in the yaw axis as seen in Fig. 15. Most of the yaw
oscillation parameter variations were tested with the SLE configuration.
In the pitch axis (Fig. 16) only a few variations in the oscillation parameter were tested with the majority of the
testing with the RLE-FT configuration. Similarly, the plunge oscillations were only conducted at a few test
conditions and are not presented graphically.
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V, m/s qmaxC re f/2V
Figure 16. Pitch oscillation frequency, velocity and amplitude test points.
7
American Institute of Aeronautics and Astronautics
All of the pitch and plunge oscillation tests were conducted in the DNW-NWB tunnel. Similarly, all of the roll
oscillation tests were done in NASA 14x22-foot tunnel. The only dynamic test conditions replicated in both tunnels
were the yaw oscillations at 1Hz and 5° amplitude at AoA’s of 10° and 14° with the RLE-FT, and at AoA’s of 10°
and 15° with the SLE.
III. Data Processing
The force and moment balance signals were sampled at 600 Hz in the DNW-NWB tests and 300 Hz in NASA
14x22-foot test. All balance signals were passed through a low-pass filter with a cut-off frequency of 5 Hz.
There were no corrections for wall or blockage effects applied to the forced oscillation time history data. The
DNW-NWB static data is available both with and without wall and blockage corrections. For comparison purposes,
all of the data presented in this paper (both static and dynamic) are without wall or blockage corrections. The
blockage from the SACCON model at 30° AoA is 4.0% in the DNW-NWB tunnel and 1.3% in the NASA 14x22-
foot tunnel.
The NASA static and forced oscillation data were corrected for sting bending. The model attitude in the DNW-
NWB tests was measured optically so no sting bending correction was required.
H. Pressure lag corrections
All pressure tube connections between the model pressure taps and ESP modules are of the same length and
diameter to assure the same time dependent behavior for each pressure tap during the unsteady pressure
measurements. The signals of the ESP modules have been corrected for attenuation and phase shift according to
Nyland et al 12 . The ten additional unsteady pressure transducers provide measurements without the pneumatic
attenuation and phase shift. However, comparisons of the unsteady pressure transducer measurements with the
Nyland corrected pressure tap measurements have not yet been completed.
I. One Cycle Averaging
As was previously noted the dynamic data runs
consisted of at least 30 seconds of multi-cycle measured
forces, moments, model position and pressures. Each
multi-cycle sinusoidal data run was later condensed to a
one cycle average loop with standard deviation about
fixed oscillation phase angle values. Figure 17 shows an
example of a 1Hz pitch oscillation pitching moment data
set for the RLE-FT configuration. The nominal AoA is
20° with a pitch amplitude of 5°. Also shown in the figure
are the static data and the 1 -cycle average with standard
deviation bars. The coefficient values at the nominal
value crossing points are of particular interest. At these
points in the oscillation cycle the rotational acceleration
is zero and the rotation rate is the maximum and
minimum. Of note is the growth in the standard deviation
of the oscillation loop in the higher AoA region of
unsteady aerodynamics. An objective of AVT-161 is to
assess the ability of current CFD methods to model the
physics required to replicate these 1 -cycle averaged
loops.
SACCON SLE
a 0 “ 20° • Static
f = 1 Hz 30s Dynamic Data
A = 5° 1 -Cycle Average & Std Dev
V = 50 m/s
Figure 17. SACCON example of pitch oscillation
1 -cycle average pitching moment coefficient verses
AoA.
J. Dynamic Derivative Analysis
Beyond the tunnel corrections and 1 -cycle averaging some additional data processing was conducted to assess
the classic dynamic derivatives from the forced oscillation data. The sinusoidal roll, pitch and yaw oscillations
include both body and velocity vector rotation rates. For example, the pitch oscillation motion has both pitch rate (< q )
and AoA rate (oc). From this type of motion a combined derivative is generally derived, such as:
Cm„ — f it? „ + C
( 6 )
American Institute of Aeronautics and Astronautics
The rolling and yawing oscillations produce combined derivates which include sideslip rate terms.
c ip = c ip + Cl. sin a (7)
^n r ^n r Cn^ COS Cl (8)
To split the combined derivative into the individual terms requires motions with only one of the rotation rates such
as plunging which has no pitch rate.
One method of deriving the combined dynamic derivatives from the oscillation data, often referred to as the
“single point” method, is to compute the difference of the maximum and minimum rate coefficient values divided by
the difference in the maximum and minimum rates 13,14 . For example:
C 171 (.Q max) Cm (q m in)
2V^ q max *7min)
( 9 )
Some examples of the combined dynamic derivatives computed in this manner are presented the following
section.
IV. Results and Discussion
The breadth of dynamic data collected during the three wind tunnel test entries exceeds what can be fully
examined within the scope of this paper. This paper will highlight the observed trends and provide selected
examples for each of the oscillation axes. The number and size of the remaining data figures prohibits them from
neatly merging with the text and are presented in the appendix.
K. Roll oscillation data
All of the roll oscillation data were collected during the NASA 14x22-foot Tunnel test with an aft sting mount,
as previously shown in Fig. 12. An example of the roll oscillation dynamic effects on the 1 -cycle averaged lateral-
directional force and moment coefficients are shown in Fig. 18 for RLE-FT configuration at AoA values from 0° to
20° in 5° increments. Up through 10° AoA the 1 -cycle averaged loops are very elliptical with minimal deviation.
Above 10° the loops are more irregular and asymmetric with increasing standard deviation. As previously noted in
Fig. 17, this deviation increase is indicative of increased unsteady flow behavior at higher AoA. The lateral
asymmetry and the vertical offset of the ellipse centroid are indicative of small asymmetries in the model geometry
and potential flow angularity. These geometric asymmetries will not likely be included in the CFD predictions for
which this data is being generated. To facilitate comparisons of the roll and yaw oscillation dynamic data with the
CFD predictions the vertical offsets at zero roll or yaw angle axis are subtracted from the dynamic data and the data
is presented as the change (A) in the forces and moments, as shown in Fig. 19. The vertical offset is computed as the
average of the two zero axis crossing values.
Figure 20 shows the roll oscillation coefficient loops for the SLE configuration at the same AoA increments,
frequency, velocity and amplitude as the RLE-FT data shown in Fig. 19. The rolling moment coefficient loops for
both leading edge configurations show similar trends with AoA. They are all elliptical and counterclockwise with
the 20° AoA loop having a significant change in slope. The elliptical shape of the coefficient loop is significant in
that it is easily modeled with the classic linear flight mechanics coefficient model.
Ci =
— ! —=C l J + C l .—+C l —+c l —
qSb/2 L p 2V L P 2V Lr 2V
( 10 )
9
American Institute of Aeronautics and Astronautics
The slope of the ellipse major axis is proportional to the C t term and the direction of the oscillation loop indicates
the sign of the damping derivative C tp . Clockwise oscillation loops represent positive derivatives (propelling),
counter-clockwise loops are negative (damping).
The yawing and side-force coefficient loops are more irregular and asymmetric and show the biggest difference
between the leading edge configurations, particularly at 15° AoA. In this AoA region there is significant change in
the forces and moments for both leading edge configurations. This will be more apparent in the pitch oscillation data
presented in the next section. Most of the remaining examples will be from this 14° to 15° AoA region.
The velocity effect was explored with the RLE-FT configuration at four velocities with the same reduced
frequency, as previously illustrated in Fig. 14. Figure 21 shows the lateral-directional coefficient loops at 14° AoA.
The lower velocities showed greater lateral asymmetry and standard deviation. At the higher velocities the loops
were more elliptical and consistent. Note however that the zero axis crossing values are nearly constant for all
velocities.
The effect of frequency at a fixed velocity is shown in Fig. 22. The higher frequency increases the vertical
thickness of the coefficient loops but maintains the general shape. The rolling moment loops remain elliptical
whereas the yawing moment and side- force loops are asymmetric.
The effect of oscillation amplitude at a fixed reduced frequency is shown in Fig. 23. The increased oscillation
amplitude tends to magnify the rolling moment loops. The yawing and side-force show not only an increase in the
size of the loops but a magnification of the asymmetry as well with the oscillation loop overlapping. The loop
overlapping is indicative of a sign reversal in the damping characteristics which may be due to flow separation at the
increased rotation rates of the higher amplitudes.
The effect of all the motion variables (frequency, amplitude and velocity) on the roll axis dynamic derivatives is
illustrated in Fig. 24 which shows the coefficient values at the zero roll angle axis crossing (peak rate, zero
acceleration) with their corresponding peak reduced rate values. The crossing pair values shown in the figure are for
all the frequency and amplitudes tested at the given velocity. Recall from the Dynamic Derivative Analysis section
that the dynamic derivative is the slope of the line through the crossing pair values. The figure shows the roll
damping derivative to be negative and invariant over a wide amplitude, frequency and velocity range. The side-force
and yawing moment show a slight shift in slope or derivative value with velocity and at large reduced rates.
Figure 25 shows all the combined roll derivative values for both leading edge configurations. The dynamic
derivatives for both leading edge configurations are roughly the same up through 10° AoA. Above 10° there is
considerable variation in the derivative values. These large variations highlight the limits of this simple linear model
to capture the non-linear character of the forces and moments at these higher angles.
L. Pitch oscillation data
All of the pitch oscillation data were collected in the DNW-NWB low-speed tunnel as listed in Table 1. The
pitch dynamic effects on the longitudinal forces and moments of the SACCON RFE-FT configuration are shown in
Figs. 26 and 27 for 1 and 3Hz oscillation frequencies, respectively. The 5° amplitude oscillations are shown about
nominal angle of attacks of 5°, 10°, 15° and 20°. The dynamic damping effect is seen as the difference between the
static and dynamic measurements. The offset between the axial force static and dynamic data is currently
unexplained and under review. The dynamic data should encompass the static data. The pitching moment showed a
larger dynamic effect than either the normal or axial force coefficients. Both force coefficients showed very little
dynamic effect below 15° AoA. In the higher AoA range near flow separation the lower frequency (1 Hz) resulted in
a more non-linear behavior than the higher 3 Hz data. This is presumed to be due to the flow dynamics having
sufficient time to transition between states at the lower frequencies. At the higher frequencies the flow does not have
time to transition resulting in a more linear behavior. The greater CFD challenge is in capturing the lower frequency
non-linear dynamic effect.
The effect of the SEE on the pitch dynamics is shown in Fig. 28 for the 1Hz pitch oscillation runs. The dynamic
effects are similar to the RFE-FT results shown in Fig. 27. The SEE axial data did not have the large offset between
the static and dynamic data seen with the RFE-FT configuration. The static data of both leading edge configurations
show a large pitch and axial force change near 15° AoA.
The dynamic testing in pitch axis did not cover as large a variation in test conditions as the roll and yaw axes, as
previously illustrated in Figs. 14-16. The effect of frequency was just reviewed and the effects of velocity and
amplitude were not explored.
Figure 29 shows the pitch axis combined dynamic derivatives for all the conditions tested. As with the roll axis
the leading edge configuration did not show a significant difference in the dynamic derivative values except at the
higher AoA where there is large dispersion in the values. The normal and axial force derivatives have very small
10
American Institute of Aeronautics and Astronautics
values up through 10° AoA with a large change in value above 15°. This characteristic was also noted in the 1 -cycle
average loops presented in Figs. 26-28. The pitching moment however showed some damping value even at the
lowest AoA values.
M. Plunging data
A relatively small set of plunging data was collected compared to the rotational oscillation testing as evidenced
in the test matrix Tables 1-4. The plunging amplitude was 0.05 meters and the free-stream velocity was 50 m/s. This
resulted in an AoA change of ±0.36° and ±0.90° for the 1.0 Hz and 2.5 Hz oscillation frequencies, respectively. An
example of the resultant longitudinal force and moment loops is shown in Fig. 30 for the 2.5 Hz plunging of the SLE
configuration.
The longitudinal AoA rate derivatives from the plunging data are shown in Fig. 31. There is not must difference
between the leading edge configurations and considerable scatter in the derivative values at 20° AoA.
N. Yaw oscillation data
An example of the yaw oscillation effects on the lateral-directional force and moment coefficients about a 15°
nominal AoA are shown in Fig. 32 for the SLE configuration at three oscillation frequencies. The lack of any
vertical surfaces on the SACCON configuration results in fairly small yaw oscillation effects. As with the roll
oscillation data there are lateral asymmetries of the data indicative of small asymmetries in the model geometry and
potential flow angularity. The increased frequency increased the vertical thickness of the force and moment loops
similar the roll and pitch data.
The yaw oscillation data for the RLE-FT configuration about a 14° nominal AoA at the same test conditions are
shown in Fig. 33. In this AoA region the leading edge appears to have a significant effect on the shape of the yaw
oscillation force and moment loops. The RLE-FT rolling and yawing moment data show very little dynamic effect in
terms of the vertical thickness of the loops and are in the opposite direction from the SLE data. The resulting
dynamic derivatives will be small negative (damping) values.
An example of the yaw amplitude effect is shown in Fig. 34 for the SLE configuration at a fixed velocity and
frequency. These data are from the NASA test and show a similar loop magnification effect as seen with the roll
amplitude in Fig. 23. The lower amplitude data was more asymmetric that the higher amplitude data which also has
higher angular rates. The higher rates may be preventing the flow from transitioning between separation and
attachment states. Future examination of the pressure time history data may reveal the dynamic flow mechanics.
The only dynamic test conditions replicated in both tunnels were the yaw oscillations at 1 Hz and 5° amplitude
with the SLE configuration at AoA’s of 10° and 15°, and with the RLE-FT configuration at AoA’s of 10° and 14°.
The tunnel to tunnel comparisons for the SLE and RLE-FT configurations are shown in Figs. 35 and 36,
respectively. There is generally good agreement in the loop shapes with the largest difference occurring at the
amplitude peaks. The zero yaw angle axis crossing values are nearly the same which will yield similar dynamic
derivatives. The repeatability of the DNW-NWB data is very good with each run 1 -cycle average loop overlaying
the other.
Figure 37 shows the yaw axis combined dynamic derivatives for all the conditions tested. As with the previous
axes the leading edge configuration did not show a significant difference in the dynamic derivative values. The
derivative values were nearly zero up through 10°. As noted in the discussion of Figs. 32 and 33 the SLE had larger
derivative magnitudes at 15° AoA than the RLE-FT at 14°, which were still near zero.
V. Summary
The three wind tunnel tests of the SACCON model in the DNW-NWB low- speed tunnel and the NASA Langley
14-by-22-Foot Subsonic tunnel have provided a wealth of static and dynamic data for CFD validation. This dataset
includes not only force and moment time histories but surface pressure and off body particle image velocimetry
(PIV) measurements as well. The extent of the data precludes a full examination within the scope of this paper. This
paper has provided some examples of the dynamic force and moment data available as well as some of the observed
trends.
The most interesting and challenging dynamics for CFD to replicate occur in the AoA range between 10° and
20° for the SACCON. Below 10° AoA the statics and dynamics are very linear with little or no dynamic hysteresis.
The only significant dynamic derivative values in this low AoA range are seen in the roll axis data and the pitching
moment data in the pitch axis. These trends are as expected for a flying wing configuration. The lack of vertical
surfaces resulted in very little yaw effect in the low AoA range.
11
American Institute of Aeronautics and Astronautics
The repeatability of the DNW-NWB dynamic data was very good. The limited tunnel to tunnel comparison data
also showed good agreement in the yaw oscillation loop shapes with the largest difference occurring at the peak
yaw deflection angles.
VI. Recommendations
The extent of the data collected during the three test of the SACCON model has created a created a considerable
data analysis challenge. This paper has highlighted some of the force and moment analysis but significant work
remains to correlate this data with the dynamic surface pressure and PIV measurements. Such analysis will provide a
better understanding of the dynamic flow topology for correlation with CFD predictions.
In an effort to minimize the aerodynamic effect of the belly mount, the attachment point and resultant rotation
axes for the pitch and yaw oscillation were well aft of the moment reference point or likely center of gravity location
for the SACCON configuration. Consequently, the pitch and yaw oscillation data from these tests are not
representative of the SACCON dynamic response when rotated about the center of gravity and should not be used as
such.. This data is however valid for the intended purpose of comparing with CFD predictions.
VII. Conclusion
A significant low-speed dataset of the SACCON model has been collected for the purpose of CFD validation.
The AoA region between 10° and 20° provides the most challenging dynamic flow topology to replicate. If CFD can
replicate the 1 -cycle averaged dynamic hysteresis loops in this AoA region then the relevant flow physics have been
captured.
Acknowledgments
The authors wish to thank the German Federal Office of Defense Technology and Procurement (BWB), which
supported the work by funding the DLR project “UCAV 2010”, and the NASA Fundamental Aeronautics Program’s
Subsonic Fixed Wing Project, which provided funding for the SACCON model and NASA wind tunnel testing.
References
Schiitte, A.; Cummings, R.; Loeser, T.; and Vicroy, D.: “Integrated Computational/Experimental Approach to UCAV and
Delta-Canard Configurations Regarding Stability & Control,” 4 th Symposium on Integrating CFD and Experiments in
Aerodynamics , Von Karman Institute, Rhode- Saint-Genese, Belgium, September 14-16, 2009.
2 Cummings, R.; and Schiitte, A.: “Computational/Experimental Approach to UCAV Stability & Control Estimation:
Overview of NATO RTO AVT-161,” AIAA-20 10-4392, 28 th AIAA Applied Aerodynamics Conference , Chicago, IL, June 28 -
July 1,2010.
3 Schutte, A.; Loeser, T.; and Oehlke, M.: “Prediction of the Flow Around the X-31 Aircraft Using Two Different CFD
Methods,” AIAA-20 10-4692, 28 th AIAA Applied Aerodynamics Conference , Chicago, IL, June 28 - July 1, 2010.
4 Jirasek, A.; and Cummings, R.: “Reduced Order Modeling of X-31 Wind Tunnel Model Aerodynamic Loads,”
AIAA-20 10-4693, 28 th AIAA Applied Aerodynamics Conference , Chicago, IL, June 28 - July 1, 2010.
5 Tomac, M.; and Rizzi, A.: “Comparing & Benchmarking Engineering Methods on the Prediction of X-31 Flying Qualities,”
AIAA-20 10-4694, 28 th AIAA Applied Aerodynamics Conference , Chicago, IL, June 28 - July 1, 2010.
6 Loeser, T.; Vicroy, D.; and Schiitte, A.: “SACCON Static Wind Tunnel Tests at DNW-NWB and 14’x22’ NASA LaRC,”
AIAA-20 10-43 93, 28 th AIAA Applied Aerodynamics Conference , Chicago, IL, June 28 - July 1, 2010.
7 Bergmann, A.; Hiibner, A.-R.: “Integrated Experimental and Numerical Research on the Aerodynamics of Unsteady
Aircraft,” 3 rd International Symposium on Integrating CFD and Experiments in Aerodynamics , ASAFA, Colorado, June 20-21,
2007.
8 Gentry, Jr., G.;Quinto, P. F.; Gatlin, G. M.; Applin, Z. T.: “The Langley 14- by 220Foot Subsonic Tunnel: Description, Flow
Characteristics, and Guide for Users,” NASA TP-3008, September, 1990.
9 Bergmann, A.; Hiibner, A.-R.; Loeser, T.: “Experimental and numerical research on the aerodynamics of unsteady moving
aircraft,” Progress in Aerospace Sciences , Volume 44, Issue 2, February 2008, Pages 121-137.
10 Owens, D. B.; Brandon, J. M.; Fremaux, C. M.; Heim, E. H.; Vicroy, D. D.: “Overview of Dynamic Test Techniques for
Flight Dynamic Research at NASA LaRC,” AIAA -2006-3146, 25 th AIAA Aerodynamic Measurement Technology and Ground
Testing Conference , San Francisco, California, June 5-8, 2006.
n Schiitte, A.; Rein, M.; Hohler, G.: “Experimental and numerical aspects of simulating unsteady flows around the X-31
configuration,” 3 rd International Symposium on Integrating CFD and Experiments in Aerodynamics , ASAFA, Colorado,
June 20-21, 2007. Proc. IMechE, Part G: J. Aerospace Engineering , 2009, 223(G4) 309-321. DOI:
1 0. 1 243/09544 1 00JAERO3 87.
12 Nyland, T.W.; Englund, D.R.; Anderson, R.C.: “On The Dynamics of Short Pressure Probes: Some Design Factors
Affecting Frequency Response,” NASA TN D-6151 (1971).
12
American Institute of Aeronautics and Astronautics
13 Brandon, J. M.; Foster, J. V.: “Recent Dynamic Measurements and Considerations for Aerodynamic Modeling of Fighter
Airplane Configurations,” AIAA 98-4447, August 10-12, 1998.
14 Brandon, J. M.; Foster, J. V.; Gato, W.; and Wilbom, J. E.: “Comparison of Rolling Moment Characteristics During Roll
Oscillations for a Low and a High Aspect Ratio Configuration,” AIAA 2004-5273, AIAA Atmospheric Flight Mechanics
Conference and Exhibit, Providence, RI, August 16-19, 2004.
15 Gilliot, A.: “Static and Dynamic SACCON PIV Tests - Part I: Forward Flow Field,” AIAA-20 10-43 95, 28 th AIAA Applied
Aerodynamics Conference , Chicago, IL, June 28 - July 1, 2010.
16 Konrath, R.; Roosenboom, E.; Schroder, A.; Pallek, D.; and Otter, D.: “Static and Dynamic SACCON PIV Tests - Part II:
Aft Flow Field,” AIAA-20 10-43 96, 28 th AIAA Applied Aerodynamics Conference , Chicago, IL, June 28 - July 1, 2010.
Appendix
Table 1 - Pitch oscillation test matrix
Tunnel
Test
Run
Config.
a
deg
Amp
deg
f
Hz
V
m/s
coc
2V
Qmax
°/s
flmaxC
2V
Duration
sec
Mount
1003-1012
0
1013-1022
RLE
5
60
0.025
0.0022
0° Yaw link
2373
1023-1032
10
1604-1613
10
5
31.4
1614-1623
1
0.0026
1084-1088
15
15° Yaw link
2426
1104-1108
0.030
1079-1083
16.5
2.5
15.7
0.0013
1094-1103
1630-1639
5
31.4
0.0026
1640-1649
RLE-FT
3
50
0.090
94.2
0.0079
DNW-NWB
1650-1659
10
1
0.030
31.4
0.0026
30
1660-1669
3
0.090
94.2
0.0079
1670-1679
1
0.030
31.4
0.0026
6° pitch link
1715-1724
15
2
0.060
62.8
0.0053
1680-1689
3
0.090
94.2
0.0079
2373
1690-1699
5
1
0.030
31.4
0.0026
1725-1734
20
2
0.060
62.8
0.0053
1700-1709
3
0.090
94.2
0.0079
1099-1108
5
60
0.025
0.0022
0° Yaw link
1109-1118
10
1194-1203
SLE
15
1
31.4
1204-1213
20
50
0.030
0.0026
15° Yaw link
1214-1223
25
Table 2 - Plunging oscillation test matrix
Tunnel
Test
Run
Config.
a
Amp
f
V
coc
Aa
adot max
Duration
Mount
deg
m
Hz
m/s
2V
deg
°/s
sec
1544-1553
10
1554-1563
15
1
0.030
0.36
2.26
1564-1573
RLE-FT
20
1574-1583
10
1584-1593
15
2.5
0.075
0.90
14.14
DNW-NWB
2373
1594-1603
20
0.05
50
30
15° Yaw link
1264-1273
10
1224-1233
15
1
0.030
0.36
2.26
1244-1253
SLE
20
1274-1283
10
1234-1243
15
2.5
0.075
0.90
14.14
1254-1263
20
13
American Institute of Aeronautics and Astronautics
Table 3 - Roll oscillation test matrix
Test Run Config.
NASA 14x22 134
10 14 15 20
10 14 15 20
10 14 15 20
10 14 15
10 15 20 25
deg Hz m/s 2V
0.24 0.06
0.36 0.09
0,44 0.11
0.55 0.14
0.57 0.15
^ 0.66 0.17
0.70 18 0.18
0.85 0.22
0.86 0.22
0
5
10
14
15
20
5
10
14
15
20
10
0
5
10
14
15
20
15
5
10
14
15
20
5
0
5
10
14
15
20
10
5
10
14
15
20
15
15 20 25
15 20 25
15 20 25
0.05
cob
2V
Pmax
°/s
Pma\b
2V
0.063
7.5
0.0055
0.095
11.3
0.0083
0.116
13.8
0.0101
0.145
17.3
0.0127
0.151
17.9
0.0131
0.174
20.7
0.0152
0.185
22.0
0.0161
0.225
26.7
0.0196
0.227
27.0
0.0198
31.4
0.0231
0.264
62.8
0.0461
94.2
0.0692
0.065
14.8
0.0057
0.096
22.0
0.0084
31.4
0.0120
0.138
62.8
0.0241
94.2
0.0361
0.064
17.9
0.0056
0.097
27.0
0.0084
n 119
31.4
0.0098
u . 1 1 z
62.8
0.0196
0.006
4.7
0.0015
0.112
94.2
0.0294
0.097
2 1 A
0.0084
J 1 .4
0.0231
0.264
62.8
0.0461
94.2
0.0692
n 119
31.4
0.0098
U . 1 1 z
62.8
0.0196
0.006
4.7
0.0015
0.112
94.2
0.0294
31.4
0.0084
0.097
62.8
0.0169
94.2
0.0253
American Institute of Aeronautics and Astronautics
Table 4 - Yaw oscillation test matrix
15
American Institute of Aeronautics and Astronautics
0.016
SACCON RLE-FT
NASA T1 34, Run 15
V= 43 m/s
A = 5 °
f= 1 Hz
0
5
10
15
20
-6 - 4-2 0 2 4 6 ''■'"-6 -4 -2 0 2 4 6
r <r
Figure 18. Example of AoA effect on roll oscillation 1-cycle averaged lateral-directional force
and moment coefficients for RLE-FT configuration.
0.002
0.02
0.001
0.01
- 0.001
<J o
- 0.01
- 0.002
- 0.02
- 0.003
- 0.03
0.006
0.004
0.002
$ 0
- 0.002
- 0.004
- 0.006
SACCON RLE-FT
NASA T1 34, Run 15
V= 43 m/s
A = 5 °
f= 1 Hz
a °
0
5
10
15
20
o
<
<t>° r
Figure 19. Example of roll oscillation lateral-directional coefficient loops with static offset
removed.
16
American Institute of Aeronautics and Astronautics
0.006
0.004
0.002
i °
- 0.002
- 0.004
- 0.006
SACCON SLE
NASA T1 34, Run 50
V= 43 m/s
A = 5 °
f= 1 Hz
a °
0
5
10
15
20
0.003
0.002
0.001
- 0.001
- 0.002
- 0.003
o
<
4>° <r
Figure 20. AoA effect on roll oscillation lateral-directional force and moment coefficients for
SLE configuration.
0.006
0.004
0.002
i 0
- 0.002
- 0.004
- 0.006
SACCON RLE-FT a = 14°
cob /(2V) = 0.097
V, m/s T134
18.3 R18
35.0 R26
43.0 R30
50.0 R16
0.003
0.002
0.001
i 0
- 0.001
- 0.002
- 0.003
-6 -4 -2 0 2 4 6
O
<
0.02
0.01
0
- 0.01
- 0.02
-6 -4 -2 0 2 4 6
<t>° <T
Figure 21. Roll oscillation velocity effect at a constant reduced frequency and amplitude for
RLE-FT about 14° nominal AoA.
17
American Institute of Aeronautics and Astronautics
0.006
SACCON RLE-FT a = 14 (
V = 43.0 m/s
0.004
0.002
£ 0
- 0.002
- 0.004
- 0.006
f, Hz T134
■ 0.57 R29
■ 0.86 R30
- 1.00 R15
0.002
0.001
% 0
- 0.001
- 0.002
-6 -4 -2 0 2 4 6
<T
O
<
<T
Figure 22. Roll oscillation frequency effect for RLE-FT about 14° nominal AoA.
0.006
0.004
0.002
£ 0
- 0.002
- 0.004
- 0.006
-20 -15 -10 -5 0 5 10 15 20
SACCON SLE <x= 15°
V = 50.0 m/s
f = 1.0 Hz
A° T134
5 R51
10 R48
15 R45
0.02
- 0.02
-20 -15 -10 -5 0 5 10 15 20
<l> ♦
Figure 23. SACCON SLE roll oscillation amplitude effect about 15° nominal AoA.
18
American Institute of Aeronautics and Astronautics
0.03
SACCON RLE-FT a = 14 (
0.02
0.01
- 0.01
- 0.02
- 0.03
-e V= 18.3 m/s
-H V = 35 m/s
V = 43 m/s
■A V = 50 m/s
Figure 24. Peak roll rate coefficient values for RLE-FT configuration at 14° nominal AoA.
SACCON
O RLE-FT
□ SLE
Figure 25. Roll axis combined dynamic derivatives.
19
American Institute of Aeronautics and Astronautics
0.09
0.08
0.07
0.06
0 E 0.05
0.04
0.03
0.02
0.01
SACCON RLE-FT
DNW-NWB T2373
V = 50 m/s
f= 1.0 Hz
A =5°
Static R1714
a 0 = 5° R1635
a 0 = 1 0° R1655
a 0 = 1 5° R1675
a 0 = 20° R1695
Figure 26. RLE-FT static and dynamic pitch oscillation longitudinal force and moment
coefficients about 5°, 10°, 15° and 20° nominal AoA at 1 Hz.
a °
SACCON RLE-FT
DNW-NWB T2373
V = 50 m/s
f= 3.0 Hz
A = 5°
— • Static R1714
a 0 = 5° R1645
a 0 = 1 0° R1665
a 0 = 1 5° R1685
a 0 = 20° R1705
a °
Figure 27. RLE-FT static and dynamic pitch oscillation longitudinal force and moment
coefficients about 5°, 10°, 15° and 20° nominal AoA at 3 Hz.
20
American Institute of Aeronautics and Astronautics
1.4
- 0.2
-5 0 5 10 15 20 25 30 35
SACCON SLE
DNW-NWB T2373
V = 50 m/s
f = 1.0 Hz
A = 5°
• Static R1284
a 0 = 1 5° R1194
a 0 = 20° R1209
a 0 = 25° R1219
a a
Figure 28. SLE static and dynamic pitch oscillation longitudinal force and moment coefficients
about 15°, 20° and 25° nominal AoA at 1 Hz.
1
o
z
O
3
2
1
0
-1
-2
-3
-4
-5
50
40
30
20
10
0
-10
<3
n
u
IS
(Si
H
H
1
i
<
i
1
-5 0 5 10 15 20 25 30
SACCON
o RLE-FT
□ SLE
X
o
2
0
-2
-4
-6
-8
-10
-12
10
15
20
25
30
Figure 29. Pitch axis combined dynamic derivatives.
21
American Institute of Aeronautics and Astronautics
SACCON SLE
V = 50 m/s
f = 2.5 Hz
A = 0.05 m
Static, R1284
a 0 = 10°, R1279
a 0 = 15°, R1239
a 0 = 20°, R1259
a ° a °
Figure 30. SLE static and dynamic longitudinal force and moment coefficients for plunging
about 10°, 15° and 20° nominal AoA at 2.5 Hz.
a°
SACCON
o RLE-FT
□ SLE
a°
Figure 31. Longitudinal angle of attack rate derivatives.
22
American Institute of Aeronautics and Astronautics
0.01
0.008
0.006
0.004
0.002
- 0.002
- 0.004
- 0.006
- 0.008
- 0.01
3=r
i — ]
i
SACCON SLE a = 15°
A = 5°
V = 50 m/s
f, Hz T2373
- 1.0, R1310
- 2.0, R1320
- 3.0, R1330
Figure 32. Yaw oscillation frequency effect at a constant velocity and amplitude for SLE about
15° nominal AoA.
SACCON RLE-FT a = 14°
A= 5°
V = 50 m/s
f, Hz T2373
- 1.0, R1459
- 2.0, R1469
- 3.0, R1479
H/° vj/°
Figure 33. Yaw oscillation frequency effect at a constant velocity and amplitude for RLE-FT
about 14° nominal AoA.
23
American Institute of Aeronautics and Astronautics
0.01
0.008
0.006
0.004
0.002
0
- 0.002
- 0.004
- 0.006
- 0.008
- 0.01
T - T
T
m
-16 -12
0
v #
12
16
0.05
0.04
0.03
0.02
0.01
0
- 0.01
- 0.02
- 0.03
- 0.04
- 0.05
SACCON SLE <x= 15°
f= 1 Hz
V = 50 m/s
A° T134
5, R61
10, R64
15, R67
j
!
A
f
A
V
/
f
—
■H
r
[ ■
#
/
/
7*
A
p
naoS^
j
i
Y
/
**
in
%
s-
H
/
>
/
i*
>
y
A
-16 -12
0
V|/°
12 16
Figure 34. Yaw oscillation amplitude effect at a constant frequency and velocity for SLE about
15° AoA.
SACCON SLE a = 15°
V = 50 m/s,
f = 1 Hz
DNW-NWB R1310
DNW-NWB R1311
DNW-NWB R1312
DNW-NWB R1313
DNW-NWB R1314
NASA LaRC R61
\j / ° V)/ °
Figure 35. SACCON SLE yaw oscillation DNW to NASA test comparison about 15° nominal
AoA.
24
American Institute of Aeronautics and Astronautics
0.008
SACCON RLE-FT a = 14'
0.006
0.004
0.002
0
- 0.002
- 0.004
- 0.006
-6 -4 -2 0 2 4 6
V = 50 m/s
f= 1 Hz
DNW-NWB R1459
DNW-NWB R1460
DNW-NWB R1461
DNW-NWB R1462
DNW-NWB R1463
NASA LaRC R7
¥°
Figure 36. SACCON RLE-FT yaw oscillation DNW to NASA test comparison about 14°
nominal AoA.
Figure 37. Yaw axis combined dynamic derivatives.
25
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